The subject matter disclosed herein relates to gas turbine systems, and more particularly to a combustor liner cooling assembly.
A combustor section of a gas turbine system typically includes a combustor chamber disposed relatively adjacent a transition piece, where a hot gas passes from the combustor chamber through the transition piece to a turbine section. As firing temperatures within the combustor chamber increase and NOx allowances are reduced, meeting combustor liner life requirements becomes increasingly challenging with currently employed cooling schemes.
One region of the combustor liner requiring effective cooling includes an aft end of the combustor liner, with one common cooling method including channel cooling. Channel cooling typically includes providing a cooling flow to a channel, then subsequently expelling the cooling flow to a region of the transition piece. Unfortunately, the useful length of the channel cooling is dependent on the temperature of the air in the cooling channel, thereby often rendering ineffective cooling of significant portions of the combustor liner due to increased firing temperatures and increased compressor discharge air temperatures. Alternatively, film cooling may be employed at various locations in the combustor chamber. Film cooling typically includes providing air from a plenum between a flow sleeve and the combustor liner to provide a barrier between the hot gas and the combustor liner. Unfortunately, the benefit of the barrier lasts for a finite length and is largely dependent on the flow in the film cooled region and not the temperature of the film gas. Therefore, either singular cooling scheme often does not achieve desired cooling performance of the aft end of the combustor liner.